Turbine rotor coolant supply system

ABSTRACT

An air supply system is configured to provide cooling air with reduced heat pickup to a turbine rotor of a gas turbine engine. The system comprises a first cooling passage extending between a hollow airfoil and an internal pipe extending through the airfoil. The airfoil extends through a hot gas path. A second cooling passage extends through the internal pipe. The coolant flowing through the second cooling passage is thermally isolated from the airfoil hot surface by the flow of coolant flowing through the first cooling passage. The first and second cooling passages have a common output flow to a rotor cavity of the turbine rotor where coolant flows from the first and second cooling passages combine according to a predetermined ratio.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a continuation of U.S. patent application Ser. No.14/974,338 filed Dec. 18, 2015, the entire contents of which areincorporated by reference herein.

TECHNICAL FIELD

The application relates generally to gas turbine engines and, moreparticularly, to a coolant supply system for providing coolant to aturbine rotor, such as a low pressure turbine (LPT) rotor.

BACKGROUND OF THE ART

It is known to provide a mid-turbine frame assembly between high and lowpressure turbine (HPT and LPT) rotors to support bearings and totransfer loads radially outwardly to a core engine casing. Themid-turbine frame assembly typically comprises a mid-turbine framesupporting an annular inter-turbine duct therein. The inter-turbine ductis defined between outer and inner duct walls which are interconnectedby a plurality of radial hollow struts, thereby forming an annular hotgas path to convey the working fluid from the HPT to the LPT. Theinter-turbine duct and the hollow struts are subjected to hightemperatures and therefore cooling air is typically introduced aroundthe inter-turbine duct and into the hollow struts to cool the same. Aportion of the cooling air supplied to the mid-turbine frame may also beused to cool the LPT rotor. However, as the air travels through themid-turbine frame, the air picks up heat. As a result, the air availablefor cooling the LPT rotor is not as cool as it could be. This may have adetrimental effect on the integrity and durability of the LPT rotor.

There is thus room for improvement.

SUMMARY

In one aspect, there is provided an air supply system for providingcooling air to a turbine rotor of a gas turbine engine, the air supplysystem comprising: at least one hollow airfoil extending through a hotgas path, at least one internal pipe extending through the at least onehollow airfoil, a first cooling passage extending between the at leastone hollow airfoil and the at least one internal pipe, a second coolingpassage extending through the at least one internal pipe, the first andsecond cooling passages being fluidly linked to the turbine rotor wherecooling air flows from the first and second cooling passages combineaccording to a predetermined ratio.

In a second aspect, there is provided a gas turbine engine comprising:first and second axially spaced-apart turbine rotors mounted forrotation about an engine axis, and a mid-turbine frame disposed axiallybetween the first and second rotors, the mid-turbine frame comprising aninter-turbine duct having annular inner and outer walls and an array ofcircumferentially spaced-apart hollow airfoils extending radiallybetween the inner and outer annular walls, the inner and outer wallsdefining a hot gas path therebetween for directing hot gases from thefirst turbine rotor to the second turbine rotor, at least one internalpipe extending through at least a first one of the hollow airfoils, afirst cooling passage extending between the at least one internal pipeand the at least a first one of the hollow airfoils, a second coolingpassage extending internally through the at least one internal pipe, thefirst and second cooling passages being fluidly linked to a rotor cavityof the second rotor.

In accordance with a still further general aspect, there is provided amethod of reducing heat pick up as cooling air travels to a turbinerotor of a gas turbine engine, the method comprising: surrounding a corecooling flow with a separate annular cooling flow while the core coolingflow travels through a hollow airfoil extending through a hot gas pathof the gas turbine engine, the annular cooling flow thermally shieldingthe core flow from thermally exposed surfaces of the hollow airfoil, andcombining the core flow and the separate annular cooling flow in apredetermined ratio to provide a common output flow to the turbinerotor.

DESCRIPTION OF THE DRAWINGS

Reference is now made to the accompanying figures, in which:

FIG. 1 is a schematic cross-section view of a turbofan gas turbineengine;

FIG. 2 is an isometric view of a mid-turbine frame and a set of externalfeed pipes for feeding cooling air to the mid-turbine frame of theengine shown in FIG. 1; and

FIG. 3 is a cross-section view of the mid-turbine frame disposed betweena HP turbine assembly and a LP turbine assembly of the engine shown inFIG. 1.

DETAILED DESCRIPTION

Referring to FIG. 1, an exemplary turbofan gas turbine engine includes afan case 10, a core case 13, an air by-pass 15 between the fan case 10and the core case 13, a low pressure (LP) spool which includes a fan 14,a LP compressor 16 and a LP turbine 18 connected by a LP shaft 12, and ahigh pressure (HP) spool, which includes a HP compressor 22 and a HPturbine 24 connected by a HP shaft 20. The core casing 13 surrounds thelow and high pressure spools to define a main fluid path therethrough.In the main fluid path, there is provided a combustor 26 to generatecombustion gases to power the HP turbine 24 and the LP turbine 18. Amid-turbine frame (MTF) 28 is disposed axially between the HP turbine 24and the LP turbine 18 and supports a bearing housing 50 containing forexample #4 and #5 bearings 102 and 104 around the respective shafts 20and 12. The terms “axial” and “radial” used for various components beloware defined with respect to the main engine axis shown but not numberedin FIG. 1.

As shown in FIG. 2, a set of external feed pipes 29 may be provided tofeed a coolant to the mid-turbine frame 28. In the illustratedembodiment, the set of feed pipes 29 comprises 4 pipes circumferentiallydistributed about the mid-turbine frame 28. However, it is understoodthat any suitable number of feed pipes may be provided. The coolant maybe compressor bleed air. For instance, the feed pipes 29 may all beoperatively connected to a source of P2.8 compressor bleed air.

As shown in FIG. 3, the MTF 28 may comprise an annular outer case 30which has forward and aft mounting flanges 31, 33 at both ends withmounting holes therethrough for connection to the HP turbine case (notshown) and the LP turbine case 36. The outer case 30, the HP and the LPturbine cases may form part of the core casing 13 schematically depictedin FIG. 1. The MTF 28 may further comprise an annular inner case 38concentrically disposed within the outer case 30. A plurality of loadtransfer spokes (not shown) may extend radially between the outer case30 and the inner case 38. The inner case 38 supports the bearing housing50 (schematically shown in FIG. 1). The bearing housing 50 may be boltedor otherwise suitably mounted to the inner case 38. The loads from thebearings 102 and 104 are transferred to the core casing 13 through theMTF 28.

The MTF 28 may be further provided with an inter-turbine duct (ITD) 40for directing combustion gases to flow generally axially through the MTF28. The ITD 40 has an annular outer duct wall 42 and an annular innerduct wall 44. An annular hot gas path 46 is defined between the outerand inner duct walls 42, 44 to direct the combustion gas flow from theHP turbine 24 to the LP turbine 18. The hot gas path forms part of theengine main fluid path. An array of circumferentially spaced-aparthollow airfoils 52 may extend radially through path 46 between the outerand inner duct walls 42 and 44. The load transfer spokes (not shown) mayextend through the airfoils 52. The airfoils 52 may be provided in theform of struts having an airfoil profile to act as turbine vanes forproperly directing the combustion gases to the LP turbine 18. As shownin FIG. 3 and as discussed herein below, the airfoils 52 may beopen-ended to fluidly connect air plenums.

As depicted by the flow arrows F1, F2 in FIG. 3, a coolant supplysystem, typically an air supply system, may be integrated to the MTF 28for supplying coolant (e.g. compressor bleed air) to the MTF 28 and theLP turbine 18. As will be seen hereinafter, the air system is configuredto minimize heat pick up as the cooling air (or other suitable coolant)travels from the source to its point of application (e.g. a rotor cavityof the LP turbine 18).

According to the illustrated embodiment, the air supply system generallycomprises at least one first cooling passage P1 extending through atleast a selected one of the hollow airfoils 52 and at least one secondcooling passage P2 extending internally through an internal pipe 60 inthe at least one selected hollow airfoil 52, the first and secondcooling passages P1, P2 having a common output flow to the rotor cavityof the LP turbine 18 where cooling air flows F1, F2 combine according toa predetermined ratio. The first cooling flow F1 flowing between theinternal pipe 60 and the airfoil 52 (the annular flow surrounding theinternal pipe 60) thermally shields the second cooling flow F2 passingthrough the internal pipe 60 from the thermally exposed surfaces of theairfoil 52, thereby reducing heat pick up as the second cooling flow F2travels radially inwardly through the hot gas path 46. In this waycooler air can be provided to the rotor of the LP turbine 18.

According to one embodiment, the air supply system may comprise twointernal pipes 60 extending through respective ones of the hollowairfoils 52. However, it is understood that any suitable number ofinternal pipes may be provided. Each internal pipe 60 is bolted orotherwise suitably connected at a radially outer end thereof to an inletport 54 provided on the outer case 30. Two of the four external feedpipes 29 (FIG. 2) are operatively connected to respective inlet ports 54and, thus, the internal pipes 60. The radially inner end of eachinternal pipe 60 is floatingly engaged with the inner case 38 of themid-turbine frame 28 for delivering a flow of cooling air in a cavity 62defined in the inner case 38 of the MTF 28. As illustrated by flowarrows F2, each cavity 62 fluidly links the associated internal pipe 60to the rotor cavity of the LP turbine 18. According to the illustratedembodiment, the second air passage P2 is, thus, defined by the internalpipes 60, the associated external feed pipes 29 and the cavities 38.However, it is understood that a different combination of componentscould be used to connect the LP turbine 18 to a source of pressurizedcooling air, while minimizing heat pick up as the air travels across thehot gas path 46.

Still according to the illustrated embodiment, the remaining twoexternal feed pipes 29 are operatively connected to an annular inletplenum 64 defined between the radially outer case 30 of the mid-turbineframe 28 and the outer annular wall 42 of the inter-turbine duct 40. Theinlet plenum 64 provides for a uniform distribution of pressurizedcooling air all around the inter-turbine duct 40, thereby avoiding localair impingement on the outer duct wall 42, which could potentially leadto hot spots and durability issues. The air directed in plenum 64ensures proper cooling of the inter-turbine duct 40. As shown by flowarrows F3 in FIG. 3, a first portion of the air received in the plenum64 flows in a downstream direction through channels defined between theouter case 30 and the LPT case 36 to pressurize and provide cooling tothe latter. More particularly, this portion of the cooling air is usedto cool and pressurize the outer shroud structure of the LP turbine 18.As depicted by flow arrows F1 in FIG. 3, a major portion of the airdirected in the inlet plenum 64 however flows radially inwardly throughthe hollow airfoils 52 about the internal pipes 60. The air isdischarged from the airfoils 52 into an outlet plenum 66 defined betweenthe inner duct wall 44 and the inner case 38 of the MTF 28. The backwall of the plenum 66 may be defined by a baffle 68 extending radiallyoutwardly from the inner case 38. Openings are defined in the baffle 68to allow air to flow in a generally downstream direction from the outletplenum 66 to the LPT rotor cavity to pressurize same and provide coolingto LPT rotor drums, as depicted by flow arrow F1 in FIG. 3. According tothe illustrated embodiment, the first cooling passage P1 is, thus,formed by inlet plenum 64, the associated external feed pipes 29, thehollow airfoils 52 and the outlet plenum 66. However, it is understoodthat a different combination of components could be used to thermallyshield the second cooling passage P2 while providing secondary air topressurize and cool the MTF 28 and associated components.

As can be appreciated from FIG. 3, the output flows from the first andsecond cooling passages P1, P2 mix together upstream from the LP turbine18 to provide a common cooling flow input to the LP turbine 18. Flowmetering devices can be provided to control the flow ratio between thefirst and second cooling passages. For instance, the flow meteringdevices could take the form of orifice plates on the feed pipes 29.According to one embodiment, 35% of the total flow to the rotor cavityof the LPT 18 originates from the first cooling passage P1. Theremaining 65% is provided via the second cooling passage P2 (i.e. thoughthe internal pipes 60). This provides for a cooler feed of cooling airto the LP turbine rotor. It is understood that the 35%-65% flow split isonly particular to a given embodiment. In fact, the flow split can bealmost anything as long as the flow distribution is sufficient topressurize the MTF 28 and ensure proper cooling of the LPT 18. The flowsplit depends on the source and sink pressure, and the failure modes ofthe system.

The above description is meant to be exemplary only, and one skilled inthe art will recognize that changes may be made to the embodimentsdescribed without departing from the scope of the invention disclosed.For example, the MTF and system and the bearing housing may have adifferent structural configuration that the one described above andshown in the drawings. Also, the air supply system could be used toprovide cooling air to a turbine rotor other than a LP turbine rotor.Still other modifications which fall within the scope of the presentinvention will be apparent to those skilled in the art, in light of areview of this disclosure, and such modifications are intended to fallwithin the appended claims.

What is claimed is:
 1. An air supply system for providing cooling air toa turbine rotor of a gas turbine engine, the air supply systemcomprising: at least one hollow airfoil extending through a hot gaspath, at least one internal pipe extending through the at least onehollow airfoil, a first cooling passage extending between the at leastone hollow airfoil and the at least one internal pipe, a second coolingpassage extending through the at least one internal pipe, the first andsecond cooling passages are fluidly linked to the turbine rotor where,when in operation, cooling air flowing from the first and second coolingpassages combine according to a predetermined ratio.
 2. The air supplysystem according to claim 1, wherein the predetermined ratio is at leastin part controlled by a flow metering device operatively connected tothe first and second cooling passages.
 3. The air supply systemaccording to claim 1, wherein the second cooling passage is configuredto deliver a greater amount of cooling air into a rotor cavity of theturbine rotor than the first cooling passage.
 4. The air supply systemaccording to claim 1, wherein the first and second cooling passages areintegrated to a mid-turbine frame of the gas turbine engine.
 5. The airsupply system according to claim 4, wherein the first cooling passagecomprises an annular inlet plenum defined between a radially outer caseof the mid-turbine frame and an outer annular wall of an inter-turbineduct defining a portion of the hot gas path, the annular inlet plenumbeing fluidly connected to a source of compressor bleed air via at leastone external pipe.
 6. The air supply system according to claim 5,wherein the at least one internal pipe is operatively connected to thesource of compressor bleed air via at least one further dedicatedexternal pipe.
 7. The air supply system according to claim 5, whereinthe first cooling passage further comprises an outlet plenum definedbetween an annular inner wall of the inter-turbine duct and a radiallyinner case of the mid turbine frame, the outlet plenum being in fluidflow communication with the rotor cavity via a baffle.
 8. The air supplysystem according to claim 7, wherein the second cooling passagecomprises at least one cavity defined in the radially inner caseradially inwardly of the outlet plenum, the at least one cavity fluidlylinking the at least one internal pipe to the rotor cavity.
 9. A gasturbine engine comprising: first and second axially spaced-apart turbinerotors mounted for rotation about an engine axis, and a mid-turbineframe disposed axially between the first and second rotors, themid-turbine frame comprising an inter-turbine duct having annular innerand outer walls and an array of circumferentially spaced-apart hollowairfoils extending radially between the annular inner and outer walls,the annular inner and outer walls defining a hot gas path therebetweenfor directing hot gases from the first turbine rotor to the secondturbine rotor, at least one internal pipe extending through at least afirst one of the hollow airfoils, a first cooling passage extendingbetween the at least one internal pipe and the at least a first one ofthe hollow airfoils, a second cooling passage extending internallythrough the at least one internal pipe, the first and second coolingpassages being in fluid communication with a rotor cavity of the secondrotor.
 10. The engine according to claim 9, further comprising a flowmetering device to control a flow split between the first and secondcooling passages.
 11. The engine according to claim 10, wherein inoperation, a majority of a total coolant flow fed into the rotor cavityis delivered via the second cooling passage.
 12. The engine according toclaim 9, wherein the first cooling passage comprises an annular inletplenum defined between a radially outer case of the mid-turbine frameand an outer annular wall of the inter-turbine duct, the annular inletplenum being operatively connected to a source of compressor bleed airvia at least one external pipe.
 13. The engine according to claim 12,wherein the at least one internal pipe is operatively connected to thesource of compressor bleed air via at least one further dedicatedexternal pipe.
 14. The engine according to claim 9, wherein the firstcooling passage further comprises an outlet plenum defined between anannular inner wall of the inter-turbine duct and a radially inner caseof the mid turbine frame, the outlet plenum being in fluid flowcommunication with the rotor cavity via a baffle.
 15. The engineaccording to claim 9, wherein the second cooling passage comprises atleast one cavity defined in the radially inner case radially inwardly ofthe outlet plenum, the at least one cavity fluidly linking the at leastone internal pipe to the rotor cavity.
 16. The engine according to claim9, wherein the first turbine rotor assembly is a high pressure turbine(HPT) rotor assembly and the second turbine rotor assembly is a lowpressure turbine (LPT) rotor assembly.
 17. The engine according to claim11, wherein at least twice as much coolant is fed via the second coolingpassage vs the first cooling passage.
 18. A method of reducing heat pickup as coolant travels to a turbine rotor of a gas turbine engine, themethod comprising: surrounding a core cooling flow with a separateannular cooling flow while the core cooling flow travels through ahollow airfoil extending through a hot gas path of the gas turbineengine, the annular cooling flow thermally shielding the core coolingflow from thermally exposed surfaces of the hollow airfoil, andcombining the core cooling flow and the separate annular cooling flow ina predetermined ratio to provide a common output flow to the turbinerotor.
 19. The method defined in claim 18, wherein surrounding a corecooling flow with a separate annular cooling flow comprises flowingcoolant in an annular space between the hollow airfoil and an internalpipe extending therethrough, the core flow flowing internally throughthe internal pipe.
 20. The method defined in claim 19, wherein combiningthe core flow and the separate annular flow in a predetermined ratiocomprises metering the amount of coolant fed to the internal pipe andthe annular space between the internal pipe and the hollow airfoil.